The invention relates generally to structure providing attitude control in an aircraft and more specifically to structure for providing yawing moments of sufficient magnitude to control the yaw in a jet-propelled aircraft flying at low speeds, at high angles of attack or in a spin.
The vertical and rudder assemblies of current aircraft generate the forces and moments necessary for directional control and stability in an aircraft by aerodynamic means wherein the deflection of the rudder imparts a sidewise acceleration to the airflow around the vertical tail to thereby create a sidewise reaction force upon the vertical tail and thus a yawing moment upon the aircraft. This side force is dependent upon the wing and fuselage flow field in which the vertical tail and rudder assembly is placed and is totally independent of thrust except for minor jet interference effects. Since conventional vertical tail and rudder assemblies are strictly aerodynamic devices, there are certain flight regimes and aircraft attitudes at which they lose their effectiveness. When an aircraft flies at low speeds or in other portions of its operating envelope of low dynamic pressure, the tail surfaces thereof may not have the capability to generate sufficient forces and moments for maintaining directional stability and control in the aircraft.
The prior art solutions used in minimizing this problem tend to degrade aircraft performance. One prior art solution to this problem is simply to increase vertical tail and rudder area to thereby increase the aerodynamic forces and moments at these flight conditions. Although this solution opens up the aircraft flight envelope, it requires that the aircraft carry the consequential increase in weight and drag throughout all other flight regimes for which the larger vertical tail is unnecessary. A second prior art solution is to limit an aircraft operating envelope such that a small vertical tail and rudder assembly is sufficiently large for effecting directional stability and control, but this solution limits aircraft maneuverability and thus runs contrary to the goal of producing highly maneuverable, high performance aircraft. Therefore, there is a definite need in the art to provide an aircraft yaw control system which retains its effectiveness at low speed flight regimes.
Another flight regime which can cause a substantial loss in directional control and stability is the operation of an aircraft at high angles of attack wherein the conventional vertical tail and rudder surfaces are washed by the fuselage and wing wake, a condition causing a loss of tail aerodynamic effectiveness, and in extreme instances, loss of aircraft control and subsequently aircraft spin. The problem is compounded by the fact that high angles of attack are required at low speeds wherein the vertical tail effectiveness has already been degraded. Prior art solutions to this problem of maintaining aircraft directional stability and control, at high angles of attack, include the attachment of ventral fins to the bottom surfaces of the aircraft, an increase in vertical tail and rudder height and mounting twin vertical tails to the sides of the aircraft at angles thereto which extend the tips of the vertical tails outwardly from the aircraft fuselage. Although these solutions improve the problem by providing tail and rudder surfaces which have some portions thereof situated outside of the fuselage and wing wake, they nonetheless degrade aircraft performance by increasing drag and weight. Furthermore, these solutions may increase tail flutter and create problems of integrating the tail surfaces into aft fuselage design. Therefore, there is a definite need in the art for an aircraft yaw control system which can retain its effectiveness at high angles of attack.
Serious problems can arise if the pilot exceeds the operational limits of an aircraft whose yaw control system relies solely upon the aerodynamic effectiveness of a conventional vertical tail and rudder assembly. Once the vertical tail and rudder surfaces lose their aerodynamic effectiveness, it is possible for the aircraft to enter a spin during which the same vertical tail and rudder must be used in the recovery of the aircraft therefrom. Even if the aerodynamic effectiveness of these surfaces could be regained during spin, a conventional vertical tail and rudder often imparts an undesirable coupling of yawing and rolling moments to the aircraft due to the rudder center of pressure being above the aircraft center of pressure. Under some conditions, this coupling may cause the spin to become more severe and preclude recovery. Because of this possibility, many aircraft carry an auxiliary spin recovery device which normally consists of a parachute attached to a long tether line. This device adds unwanted weight to the aircraft and increases drag because the parachute housing must be faired into the clean aircraft lines. Therefore, there is a definite need in the art to provide an aircraft yaw control system which retains its effectiveness throughout spin rcovery.
Accordingly, it is an object of the present invention to provide a relatively lightweight and drag-free yaw control system in a jet propelled aircraft which retains its effectiveness at low flight speeds, at high angles of attack and throughout a spin.
Another object of the present invention is to provide a yaw control system which produces minimal adverse coupling of yawing and rolling moments when used for spin recovery.
Another object of the present invention is to provide an aircraft yaw control system which is conductive to aft, fuselage-empennage integration.